Decomposed ammonia radioisotope thruster

ABSTRACT

A radioisotope heated propellant reaction control system wherein energy from the decay of plutonium-238 is used to heat and decompose ammonia propellant, and the decomposition products, nitrogen and hydrogen, are expanded through nozzles to provide desired increments of thrust. Three nozzles provide thrust levels of 10 X 10 3 to 100 X 10 3 lbf at a specific impulse of 310 seconds in a pulsing mode.

I llmted States Patent 11 1 1111 3,724,215 Neudecker et al. [4 1 Apr. 3,1973 541 DECOMPOSED AMMONIA 3,569,714 3 1971 Anderson @161 ..250l108RRADIOISOTOPE THRUSTER 3,197,959 8/1965 Keller ..60/229 3,302,042 l/1967Grover et al. .....176/39 lnventorsl Joseph Neudecker; Kenneth 3,516,4876 1970 Keiser ..176/39 p both of L05 Alamos, 3,447,321 6/1969 Romero.l76/39X Mex. 3,525,386 8/1970 Grover. l65/105X [73] ss g e e Unitedsmtes America as 3,451,641 6/1969 Leventhal ..165/105X i w h [BindSuites Primary Examiner-Clarence R. Gordon nergy ommlss'on AssistantExaminer-Robert E. Garrett [22] Filed: May 19, 1971 Attorney-Roland A.Anderson [21] Appl.No.: 144,954 [57] ABSTRACT A radioisotope heatedpropellant reaction control "60/203 system h i v gy f m the de ay ofplutonium- [58] Field 6: Search ..60/203, 225, 229, 266; 176/39; 238 andF Pmpe" 250/106 8 108 165/105 lant, and the decomposmon products,mtrogen and hydrogen, are expanded through nozzles to provide [56]References Cited des1red Increments of thrust. Three nozzles provldeUNITED STATES PATENTS Friedman et a1. ..60/203 thrust levels of 10 X 10'to 100 X 10" lb, at a specific impulse of 310 seconds in a pulsing mode.

- 4 Claims, 4 Drawing Figures PATENTEUAPR 3 I975 SHEET 1 OF 2 0 0 8 2 Nw I w S o P M o m N D 2 w T O I R w E l P M w E lw In D 2 o m o I 0 4 Mw W O O a m. m 8

TEMPERATURE (F) INVENTOR. Kenneth C. Cooper BY Joseph W Neudec/rer Fig.2

DECOMPOSED AMIVIONIA RADIOISOTOPE THRUSTER BACKGROUND OF THE INVENTIONThe invention described herein was made in the course of, or under, acontract with the US. ATOMIC ENERGY COMMISSION. It relates toradioisotope heated propellant reaction control systems and moreparticularly to a system wherein energy from the decay of a radioisotopeis used to heat and decompose ammonia propellant, and the decompositionproducts, nitrogen and hydrogen, are expanded through a nozzle toprovide desired increments of thrust.

Many communications and observational satellites require small vernierrocket engines for (l) orbital injection error correction, (2)station-keeping and repositioning, and (3) attitude acquisition andcontrol. This is particularly true for a satellite placed in an earthsynchronous or geosynchronous orbit which is a circular orbit in theequatorial plane with an orbital period of one sidereal day. In such anorbit the satellite will ideally remain fixed in space, relative to anobserver on the earth. Satellites in these orbits typically performmissions as communications relays, navigational aids, and meteorologicaland strategic reconnaissance vehicles. Many of these missions requireextremely close pointing accuracy and precise stationkeeping for aperiod of years.

A decomposed ammonia radioisotope thruster (DART) in which aradioisotope is used to heat and decompose ammonia prior to expansionthrough a nozzle offers significant improvement in specific impulse,weight, and reliability over chemical vernier engine propulsion systemsnow in use. However, any propulsion system utilizing a radioisotopeenergy source has several drawbacks not experienced by its chemicalsystem counterparts. One is that it must meet nuclear safetyspecifications which require that the radioisotope remain completelycontained within its fuel capsule during any launch pad or otheraccident and also during atmospheric reentry and impact. Another is thatthe radioisotope is a constant and essentially nonvariable heat sourcefor satellite life spans up to years. While this is an advantage in thevacuum of space, it requires that some form of cooling be employed inatmosphere to avoid materials degradation through oxidation caused byelevated temperatures.

A significant problem in the design of DART propulsion systems has beenensuring that the radioisotope fuel capsule will survive intact thethermal and pressure stresses and temperature it would be subjected toduring atmospheric reentry. Efforts to resolve this problem heretoforehave centered on placing the entire thruster unit within a reentrymodule.

We have now found that the radioisotope fuel capsule and its associatedimpact capsule can be effectively protected against the thermal andpressure stresses and temperature of atmospheric reentry by placing themwithin a reentry ablation capsule. All other parts of the DART unit, asfor example, the thruster nozzles, insulation, and ammonia propellantheat transfer system, are outside the reentry capsule and would beallowed to burn up on reentry. The reentry capsule is required to bemade of an ablative material having good thermal conductivity sinceduring the operational life of the DART unit heat must be effectivelyand efficiently transferred from the radioisotope through the reentrycapsule to the ammonia propellant heat transfer system. We have foundthat ATJS graphite is an excellent ablative material for this purpose.

We have further found that by using heat pipes inserted into the reentryablative capsule to transfer excess heat to the atmosphere, thetemperature within the DART unit can be kept under 200 C, thuspreventing refractory structural materials from oxidizing before thesatellite is inserted into orbit. The heat pipes represent a trulypassive coolant system since they require no auxiliary pumping ormonitoring apparatus.

It is therefore an object of this invention to provide an improvedvernier engine for precise satellite control and stabilization. Anotherobject is to provide an improved vemier engine comprising a decomposedammonia radioisotope thruster (DART) unit. A further object is toprovide a DART unit in which the radioisotope fuel capsule will remainintact during a launch pad or other accident and during atmosphericreentry and impact. Still another object is to provide a method by whicha DART unit can be passively cooled.

BRIEF DESCRIPTION OF THE DRAWINGS These and other objects of thisinvention will be apparent from the following description read inconjunction with the accompanying drawings wherein:

FIG. 1 is a schematic of a DART propulsion system.

FIG. 2 shows the theoretical specific impulse of ammonia as a functionof temperature.

FIG. 3 is a cross-sectional view of the preferred embodiment of the DARTunit.

FIG. 4 is a cross-sectional view taken as indicated by the line 4-4 inFIG. 3.

The operation of a DART propulsion system is readily described withreference to the schematic of FIG. 1. Liquid ammonia 4 is introducedinto a propellant storage tank 11 through a fill valve 3 and storedunder its own equilibrium, vapor pressure. Located within tank 11 may bea heater 5 to supply adequate heat of vaporization for the ammonia andto maintain a minimum tank pressure should ambient temperature fallbelow approximately 50 F. By means of pressure transducers 6 and apressure control valve 1, pressure within tank 11 is prevented frombecoming excessive. Should the pressure become too high, ammonia gas canbe bled from tank 11 through relief valve 2. Nominally, the temperaturewithin tank 11 will be F and the pressure will be about 200 psia. Duringoperation of the system, gaseous ammonia is introduced into aradioisotope thruster unit 8 through electrically controlled pulsinginlet valves 7. Nominally, pressure of ammonia to the inlet valves 7 isregulated at 50 psia by means of a fine pressure regulator 10. Withinradioisotope thruster unit 8 the gaseous ammonia is heated anddecomposed into hydrogen and nitrogen. The hot mixture of hydrogen andnitrogen is then allowed to expand through thrust nozzles 9 to provide adesired increment of thrust. Each of the inlet valves 7 provides anammonia flow for a particular thrust nozzle 9. The invention describedherein relates to radioisotope thruster unit 8.

It is apparent from FIG. 2 that the specific impulse that can beachieved using ammonia as a propellant is dependent on the temperatureto which the ammonia is heated. At any particular temperature above thedecomposition temperature of ammonia, a further substantial increase inspecific impulse can be achieved if the ammonia is in fact completelydecomposed to nitrogen and hydrogen and the mixture of these gasesallowed to expand through a thrust nozzle.

PREFERRED EMBODIMENT FIGS. 3 and 4 are cross-sectional views of thepreferred embodiment of this invention. The DART unit attains an exitgas temperature of 1370 C (2500 F) through the thrust nozzles andprovides thrust levels of 10 X 10 to 100 X 10' lb, at a specific impulseof 310 seconds in a pulsing mode for a total satellite lifetime of sevenyears. There are three separate thrust nozzles 12, 13, and 14 with thefollowing individual characteristics and thrust specifications:

Nozzle l2 13 I4 Diameter of Throat 0.016 inch 0.03l inch 0.049 inchThrust I mlb 50 mlb 100 mlb Maximum Pulses 5 X 5 X 10 5 X l0 NH TotalFlow [00 lb 100 lb 345 lb NI-l; Flowrate (lb/sec) 3.25 X 10 1.62 X 10'3.25 X 10" Maximum Orr-Time 200 msec 200 msec 200 msec Minimum On-Timemsec 20 msec 20 msec Minimum Off-Time 800 msec 2 sec 4 sec In additionto nozzles 12, 13, and 14, the other major components of the DART unitare the radioisotope heat source 15, the impact capsule 16, the ablativereentry capsule 17, the heat transfer system 18, the insulation system19, and the mounting can 20.

Plutonium-238, which is an alpha emitter having a half-life of 89 years,serves as the heat source 15. The plutonium is present in the form ofcermet fuel wafers 21 which consist of hot pressed molybdenum-coatedparticles of either PuO ThO solid solution or solely PuO Methods offabricating these fuel wafers are given in Los Alamos ScientificLaboratory Reports LA- 4476-MS and LA-4647-MS, available from NationalTechnical Information Service, U. S. Department of Commerce, 5285 PortRoyal Road, Springfield, Virginia 22151. A wafer 21 is approximately 2.2inches in diameter and 0.22 inch thick and provides 40 watts of thermalpower. Six wafers are deposited within fuel capsule 22 which issupported within impact capsule 16 by cushions 23 and 24. These cushionsconsist of molybdenum foamed to 35 percent of theoretical density.

Certain requirements are imposed on the materials of fuel capsule 22 andimpact capsule 16. They must be strong at elevated temperatures, have ahigh creep and impact resistance over the temperature range from ambientto l500 C, and be ductile both at ambient and elevated temperatures. Inaddition, they must not chemically react with the fuel at anytemperature below 1800 C. Molybdenum-rhenium alloys selected from therange of Mo-46 wt. percent Re to Mo-2O wt. percent Re have been found tomeet these overall requirements and hence are used as the material offuel capsule 22 and impact capsule 16. Although all Mo- Re alloysoxidize, those having a lower weight percent of rhenium tend to havelower oxidation rates, especially at elevated temperatures. Hence, alloyselection is a compromise between strengthductility values and anacceptable oxidation rate. In the preferred embodiment of thisinvention, Mo-46 wt. percent Re is used for fuel capsule 22 and impactcapsule 16. Both capsules are assembled by welding, with impact capsule16 being comprised of a female member 25 and a male member 26.

The isotope fuel being an alpha emitter results in the gradual releaseof helium gas from the fuel wafers. Over a longer period of time thishelium gas, if not provided with a path of egress, would cause pressureto build up inside thefuel capsule 22 and impact capsule 16. The fuelcapsule has a thin wall so that the helium gas pressure will rupture itquite easily. A 0.005-in. diameter vent hole 55 is provided in theimpact capsule to permit the helium gas to vent to the vacuum of space.

The ablative reentry capsule 17 is composed of a cylinder 27 and two endplugs 28 and 29 of ATJS graphite. A reentry capsule must protect theimpact capsule and the fuel against the heat generated during reentry byfriction with the earths atmosphere. Graphite is commonly used for thereentry capsule material, as it withstands high temperatures and absorbssome of the reentry heat by oxidative ablation of the surface.

Disposed immediately adjacent to but not in direct contact with graphitecylinder 27 is heat transfer system 18. This system consists of aninner, thicker cylinder 31 and an outer thinner cylinder 32. In thepreferred embodiment of the invention, cylinders 31 and 32 are made ofrhenium which is chemically inert to ammonia and its decompositionproducts, hydrogen and nitrogen, at temperatures up to 2500 C. It alsocatalyzes the decomposition of ammonia. Disposed within cylinder 31 andbetween cylinders 31 and 32 are three spiral grooves 33, 34 and 35through which ammonia flows and is decomposed. Each groove makes onecomplete spiral from an ammonia inlet to a thrust nozzle. The ammoniainlet for a particular thrust nozzle is located directly in line withthe nozzle it serves. Thus, ammonia inlet 36 connects by means of groove34 with thrust nozzle 13, while ammonia inlet 37 connects by means ofgroove 35 with thrust nozzle 14. Cylinders 31 and 32 are electron beamwelded together between the spiraled grooves, thus forming each grooveinto a channel through which ammonia and its decomposition productsflow.

For efficient operation of the DART unit it is essential that heattransfer system 18 be effectively insulated from the space environment.It is desirable that insulating system 19 function such that there is aminimum variation in temperature in heat transfer system 18 between fullthrusting with maximum heat transfer to the propellant and zero thrustso as to maintain as closely. as possible a nominal thruster operatingtemperature of 1370 C. To ensure this, insulating system 19 operatesprimarily by the principles of radiation heat transfer, in which theheat transfer varies as the fourth power of the temperature. In thepreferred embodiment of the invention, insulating system 19 is composedof an outer insulation cup 38 made of stainless steel, insulation 39,and an inner insulation cup 40 made of molybdenum. Insulation 39consists of layers of molybdenum foil, 0.0005-in. thick, each layerseparated from its neighboring layers by zirconia present either in theform of particles or as woven zirconia cloth.

The interdiffusion of materials becomes a problem at the elevatedtemperatures present in the DART unit. Therefore, some differentmaterials must be separated from each other by diffusion barriers. Thesediffusion barriers are incorporated into the support structure; hence,they effectively serve dual purposes. The support structure separatingMo-Re impact capsule 16 from graphite ablative reentry capsule 17consists of tantalum disks 41 and 42, which carburize to tantalumcarbide at elevated temperatures, and zirconium carbide support cones 43and 44. Rhenium heat transfer system 18 is separated from graphiteablative reentry capsule 17 by tantalum wire 45 wrapped around cylinder27. This provides a gap 56 about mils wide between heat transfer system18 and cylinder 27. At operating temperatures, wire 45 becomes tantalumcarbide which acts as a diffusion barrier. Ablative reentry capsule 17is supported within insulation system 19 by tantalum carbide cones 46and 47, also good diffusion barriers. Mounting can 20 is attached toinsulating system 19 by means of zirconia bushings 48 and 49 and held inproper orientation to insulating system 19 by the disk spring action ofthe flat ends of mounting can 20. The outer insulation cup 38 is held inaxial position by leaf springs 50 and 51.

The DART unit, being insulated, attains an internal temperature ofapproximately 1400 C during operation in the vacuum of space. However,refractory structural materials such as rhenium and molybdenum begin tosignificantly oxidize at temperatures above 300 C. To prevent this, itis necessary during all time prior to launch to either isolate the unitfrom oxygen or cool it to temperatures below 300 C. The latter approachis the most feasible. In the preferred embodiment of this invention, twoheat pipes using water as the working fluid provide the necessarycooling. Such heat pipes 60 are well described in the literature. Theheat pipes are placed in wells in cylinder 27 and extend throughinsulation system 19 and mounting can 20. During storage of the DARTunit or prior to launch, these heat pipes radiate and convect heat tothe atmosphere. Before launch they are removed and the vehicle islaunched expeditiously so that the unit is placed in orbit prior toattaining a temperature of 300 C. Alternatively to the use of heatpipes, cold nitrogen may be circulated through heat transfer system 18to cool the unit before launch.

It will be apparent to one of ordinary skill in the art that what hasbeen disclosed is a vernier rocket engine system having low electricalpower requirements, relatively high specific impulse, high reliabilityof operapulsing capabilities, and a weight less than that of comparablechemically fueled systems. It will be further apparent that the DARTunit herein disclosed is but a specific embodiment of a more generalfamily of decomposed ammonia radioisotope thruster vernier rocketengines, and that the design and operating parameters of the unitdisclosed by example herein may readily be varied to meet a wide varietyof orbital station keeping and attitude acquisition and controlrequirements.

What we claim is: 1. A radioisotope heated propellant reaction controlsystem wherein energy from the decay of plutonium- 238 is used to heatand decompose ammonia propellant, and the decomposition products,nitrogen and hydrogen, are expanded through nozzles to provide desiredincrements of thrust, comprising in combination,

a plutonium-238 fuel,

a fuel capsule for containing said fuel,

an impact capsule for containing said fuel capsule,

a graphite ablative reentry capsule,

a heat transfer system containing a plurality of channels whereinammonia is heated and decomposed to nitrogen and hydrogen,

a plurality of ammonia inlet tubes connected to said heat transfersystem, each of said tubes connected to the entrance end of a particularchannel in said heat transfer system,

means for controlling the flow of ammonia through said inlet tubes,

a plurality of thrust nozzles, each of said nozzles connected to theexit end of a particular channel in said heat transfer system,

an insulating system disposed adjacent to the outer periphery of saidheat transfer system, with said thrust nozzles extending through saidinsulating system, and

a mounting can,

said ablative reentry capsule encompassing said impact capsule anddisposed between said impact capsule and said heat transfer system.

2. The reaction control system of claim 1 wherein said plutonium-238fuel comprises wafers of hotpressed molybdenum-coated PuO particles.

3. Thereaction control system of claim 1 wherein said plutonium-238 fuelcomprises wafers of hotpressed molybdenum-coated solid solution PuO Th0cermet particles.

4. The reaction control system of claim 1 containing a plurality ofwater heat pipes inserted into wells in said ablative reentry capsule tocool said system below 200 C, said heat pipes extending through theinsulation system and the mounting can and being removable tion, alifetime in excess of seven years, a propellant from Said ablativereentry capsule before launchfree of troublesome handling problems,excellent

2. The reaction control system of claim 1 wherein said plutonium-238fuel comprises wafers of hot-pressed molybdenum-coated PuO2 particles.3. The reaction control system of claim 1 wherein said plutonium-238fuel comprises wafers of hot-pressed molybdenum-coated solid solutionPuO2-ThO2 cermet particles.
 4. The reaction control system of claim 1containing a plurality of water heat pipes inserted into wells in saidablative reentry capsule to cool said system below 200* C, said heatpipes extending through the insulation system and the mounting can andbeing removable from said ablative reentry capsule before launch.